1. Field of the Invention
The present invention is directed to a clamping connection assembly for retaining separable component parts of a spacecraft and more particularly, to a V-clamp connection assembly capable of preventing relative rotational movement and transmitting shear and torsional forces without changing the axial load created by a tension strap used to retain the separable components of a spacecraft.
2. Description of the Prior Art
Connection assemblies, for example, those that retain separable parts of a spacecraft, provide important structural links in the primary load path of a spacecraft during both boost and ascent flight and also, if necessary, during re-entry. For example, satellites frequently jettison a portion of their structure, such as a propulsion motor, as it is lifted into a permanent orbit. As can be readily appreciated, spacecraft connection assemblies must be designed not only for low weight, but also for extremely high reliability due to the high cost of the spacecraft launch and the general inaccessibility to rectify any errors.
As this particular field of art has advanced from the 1950s to the present day, there has been a recognition of the design requirement of minimizing the component parts that can fail. Even today with the realization of the launching of satellites and spacecraft from the Space Shuttle, the high cost involved in employing a human being in space still requires a connection assembly design that will be automatic, without human intervention and with an extremely high reliability. Likewise the design criteria of minimizing weight remains as important as ever giving the high cost of transmitting the payload of a spacecraft into outer space. Designers in this field have recognized the desirability of using tension band clamping assemblies, instead of exploding connecting assemblies, to hold together separable components of a spacecraft and thereby minimizing any pyro-shock that could adversely vibrate the payload of the spacecraft, such as electronic components.
In the 1960s, a connection assembly design referred to as the Marman Band was proposed which included a plurality of V-retainers or wedged shaped shoe members that provided a V-clamping to flanges positioned on the exterior surfaces of separable component parts of the spacecraft. A pair of separable tension bands were joined together by a pair of spaced pyrotechnic tension bolts that could adjustable vary the tension in the straps that held the V-band coupling members. The tension straps were utilized to create inwardly radial forces on the V-shaped retainers which would in turn, wedge the flanges of cylindrical housing structures of separable component parts of the spacecraft in an axial direction. This V-band coupling was basically configured to transfer axial loads and bending moments across the resulting structural joint of the connection assembly. Small shear torque loads were transferred across the joint by relying on friction that would exist between the flange surfaces and the retainer wedges. If the spacecraft was to be subject to considerable forces that would produce large shear torque loads, either keys were inserted between the flanges or the friction load that was applied during the loading of the tension straps, had to be carefully controlled to take into consideration the magnitude of the shear torque forces that were to be expected.
Thus prior to the advent of the Space Shuttle, the primary connection assembly relied upon for separable spacecraft component parts were V-band couplings that were encircled with tension straps to maintain an axial loading between the spaceship component parts. Usually the friction of the band and the axial loading friction on the flanges or cylindrical structures of the spacecraft component parts were relied upon to handle any shear and torque forces.
The cause of the shear and torque forces resulted from many separate sources such as transitory forces due to the cyclic nature of a rocket burn particularly in solid fuel rockets, variances in the center of gravity of the fuel container during the burn-off, changes in nozzle direction to control the rocket, changes in trajectory of the rocket to achieve the desired orbit path, flexing of the rocket itself during flight, vibrations, etc. Usually however, the ratio of the longitudinal forces generated during flight to any transverse loading was about 10:1.
However, a significant increase in shear and torque forces occurred when spacecraft were to be launched from the bay of the Space Shuttle. As can be seen in FIG. 4, the spacecraft payload was positioned offset from the principle rocket thrust from the rockets 102 and this position can produce force ratios that approached a 1:1 ratio between the axial force loading and the transverse loading. Additionally, the spacecraft was usually mounted to be cantilevered from releasable restrains that permitted additional transverse vibration to be created, since the minimum number of restraining points provided the least number of possible failure points. Additionally, the Space Shuttle 100 also required designing any clamp connection assembly to take into consideration the possibility of re-entry landing if necessary and its associated high transverse loading.
When the magnitude of the transverse loading was appreciated, attempts were made to add a spline between the flanges of the separable component parts of the spacecraft to resolve the problem. It was recognized that in increasing the capacity of transmitting the shear and torsional forces that an increase in the possibility of seizing between the separable components parts would occur. It was also recognized that metallic parts that rub together had a higher probability of seizing or even welding together in outerspace than doing so on earth.
Additionally, the use of the Space Shuttle further provided an ability to increase the diameter of spacecraft as compared to rocket launched spacecraft and with the larger diameters, there was a greater need to transmit the shear and torsional forces and thereby prevent any relative rotational movement between the separable spacecraft component parts.
The increased shear carrying requirement for shuttle launched spacecraft is derived as follows:
Transverse force balance ##EQU1## where S=mg.sub.T Transverse force
P.sub.ax =mg.sub.A Axial force PA1 m=Spacecraft means PA1 D=clamp diameter PA1 H=spacecraft center of weight above connection clamp PA1 g.sub.T =transverse acceleration PA1 g.sub.A =axial acceleration PA1 f=coefficient of friction at clamp PA1 (shear carrying requirement)
Substitution yields ##EQU2##
This equation (2) is best evaluated by inserting typical values for a shuttle and expendable rocket launched spacecraft as listed in Table 1.
TABLE 1 ______________________________________ Configuration Load factors parameters Shear factor g.sub.A g.sub.T H D f ______________________________________ Shuttle 4 6 40 90 .41 Expendable 16 1.5 30 36 .07 Rocket ______________________________________
The shear factor for a shuttle launched spacecraft far exceeds that of an expendable rocket and a factor of 0.4 can not reliably be provided by friction. Thus, a mechanical shear bearing member is needed.
An additional limitation in the design of connection assemblies was the clear recognition that the tension band design had reached a fairly developed and reliable state of art and it was necessary in resolving the problems presented by this increase in shear and torsional forces that no additional load would be asserted on the tension band that would require further redesign.
One of the proposed prior art solutions to this problem is illustrated in FIG. 7, wherein a connector assembly 200 utilizes a series of aluminum wedge blocks 202 that are forced against the joint flange members 206 and 208 by a clamp tension band 204. The internal surface of the flange member 206 has a series of conical detents of approximately 60 in number, that are spaced about the diameter of the flange, for receiving shear cones 214 that are adjustably mounted on threaded shafts 212 journalled within threaded ports 210 in the lower flange member 208. Thus, the individual shear cones could be adjusted to eliminate any possible gap in the detent on the flange member 206. It is necessary, however, to ensure that no seizing or binding occurs and therefore testing on appropriate and expensive test jigs is required to determine if the shear cones are not binding on the separable spacecraft component parts, for example, by physically separating the component parts during spacecraft testing. As can be readily appreciated, the slanting surfaces of the cones which were positioned at conical surfaces that subscribed at an acute angle of 15.degree. to 20.degree., provided in effect, camming surfaces and any shear load application as well as misalignment would produce corresponding counter axial forces that would increase the radial force load on the tension band.
Thus there is still a need in this aerospace field to improve the clamping connection assembly that is used to hold two separable component parts of a spacecraft together.